EP3361158B1 - Brennkammer für eine gasturbine - Google Patents

Brennkammer für eine gasturbine Download PDF

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Publication number
EP3361158B1
EP3361158B1 EP18156681.1A EP18156681A EP3361158B1 EP 3361158 B1 EP3361158 B1 EP 3361158B1 EP 18156681 A EP18156681 A EP 18156681A EP 3361158 B1 EP3361158 B1 EP 3361158B1
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EP
European Patent Office
Prior art keywords
combustor
liner
rail
furrow
liner panel
Prior art date
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Active
Application number
EP18156681.1A
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English (en)
French (fr)
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EP3361158A1 (de
Inventor
Steven W. Burd
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RTX Corp
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United Technologies Corp
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/002—Wall structures
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R3/00—Continuous combustion chambers using liquid or gaseous fuel
    • F23R3/007—Continuous combustion chambers using liquid or gaseous fuel constructed mainly of ceramic components
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03042—Film cooled combustion chamber walls or domes
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F23—COMBUSTION APPARATUS; COMBUSTION PROCESSES
    • F23R—GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS
    • F23R2900/00—Special features of, or arrangements for continuous combustion chambers; Combustion processes therefor
    • F23R2900/03044—Impingement cooled combustion chamber walls or subassemblies
    • Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
    • Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
    • Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
    • Y02T50/00—Aeronautics or air transport
    • Y02T50/60—Efficient propulsion technologies, e.g. for aircraft

Definitions

  • the present disclosure relates to a gas turbine engine and, more particularly, to a combustor section therefor.
  • Gas turbine engines such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. Relatively high temperatures are observed in the combustor section such that cooling airflow is typically provided to meet desired service life requirements.
  • the combustor section typically includes a combustion chamber formed by an inner and outer wall assembly.
  • Each wall assembly may include a support shell lined with heat shields often referred to as liner panels.
  • the combustor includes liner panels with a hot side exposed to the gas path and an opposite, or cold side, that has features such as cast in threaded studs to mount the liner panel and a full perimeter rail that contact the inner surface of the combustor liner support shells.
  • the wall assemblies are segmented to accommodate growth of the panels in operation and for other considerations.
  • Combustor panels typically have a quadrilateral projection (i.e. rectangular or trapezoid) when viewed from the hot surface.
  • the panels have a straight edge that forms the front or upstream edge of the panel and a second straight edge that forms the back or downstream edge of the combustor.
  • the panels also have side edges that are linear in profile.
  • the liner panels extend over an arc in a conical or cylindrical fashion in a plane and terminate in regions where the combustor geometry transitions, diverges, or converges. This may contribute to durability and flow path concerns where forward and aft panels merge or form interfaces. These areas can be prone to steps between panels, dead regions, cooling challenges and adverse local aerodynamics.
  • US 2016/252249 A1 discloses a prior art combustor as set forth in the preamble of claim 1.
  • a combustor for a gas turbine engine according to claim 1.
  • a further embodiment of the present invention may include that the rail is at least in partial contact with a surface formed by the furrow.
  • a further embodiment of the present invention may include that the furrow forms a concave surface that faces a liner panel.
  • a further embodiment of the present invention may include that the furrow includes at least one cooling impingement passage therethrough, the cooling impingement passage directed toward the rail.
  • a further embodiment of the present invention may include that the second liner panel is an aft liner panel.
  • a further embodiment of the present invention may include that the furrow includes a displaced surface parallel to a combustor liner support shell surface that faces the first and second liner panels, a first surface that extends between the displaced surface and the surface, and a second surface that extends between the displaced surface and the surface.
  • a further embodiment of the present invention may include that the first surface and the second surface are angled with respect to the surface and the displaced surface.
  • FIG. 1 schematically illustrates a gas turbine engine 20.
  • the gas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28.
  • Alternative engine architectures 200 might include an augmentor section among other systems or features.
  • the fan section 22 drives air along a bypass flowpath and into the compressor section 24.
  • the compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26, which then expands and directs the air through the turbine section 28.
  • turbofan in the disclosed non-limiting embodiment, it should be appreciated that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans wherein an intermediate spool includes an intermediate pressure compressor ("IPC") between a Low Pressure Compressor (“LPC”) and a High Pressure Compressor (“HPC”), and an intermediate pressure turbine (“IPT”) between the high pressure turbine (“HPT”) and the Low pressure Turbine (“LPT”).
  • IPC intermediate pressure compressor
  • LPC Low Pressure Compressor
  • HPC High Pressure Compressor
  • IPT intermediate pressure turbine
  • the engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine static structure 36 via several bearing structures 38.
  • the low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor (“LPC”) 44 and a low pressure turbine (“LPT”) 46.
  • the inner shaft 40 drives the fan 42 directly or through a geared architecture 48 to drive the fan 42 at a lower speed than the low spool 30.
  • An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
  • the high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor (“HPC”) 52 and high pressure turbine (“HPT”) 54.
  • a combustor 56 is arranged between the HPC 52 and the HPT 54.
  • the inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
  • Core airflow is compressed by the LPC 44, then the HPC 52, mixed with the fuel and burned in the combustor 56, then expanded over the HPT 54 and the LPT 46.
  • the LPT 46 and HPT 54 rotationally drive the respective low spool 30 and high spool 32 in response to the expansion.
  • the main engine shafts 40, 50 are supported at a plurality of points by bearing systems 38 within the static structure 36.
  • the gas turbine engine 20 is a high-bypass geared aircraft engine.
  • the gas turbine engine 20 bypass ratio is greater than about six (6:1).
  • the geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system.
  • the example epicyclic gear train has a gear reduction ratio of greater than about 2.3, and in another example is greater than about 2.5:1.
  • the geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the LPC 44 and LPT 46 and render increased pressure in a fewer number of stages.
  • a pressure ratio associated with the LPT 46 is pressure measured prior to the inlet of the LPT 46 as related to the pressure at the outlet of the LPT 46 prior to an exhaust nozzle of the gas turbine engine 20.
  • the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1)
  • the fan diameter is significantly larger than that of the LPC 44
  • the LPT 46 has a pressure ratio that is greater than about five (5:1). It should be appreciated, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans.
  • a significant amount of thrust is provided by the bypass flow path due to the high bypass ratio.
  • the fan section 22 of the gas turbine engine 20 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet (10668m). This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC).
  • TSFC Thrust Specific Fuel Consumption
  • Fan Pressure Ratio is the pressure ratio across a blade of the fan section 22 without the use of a Fan Exit Guide Vane system.
  • the low Fan Pressure Ratio according to one non-limiting embodiment of the example gas turbine engine 20 is less than 1.45.
  • Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of ("Tram" / 518.7) 0.5 .
  • the Low Corrected Fan Tip Speed according to one non-limiting embodiment of the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
  • the combustor section 26 generally includes a combustor 56 with an outer combustor wall assembly 60, an inner combustor wall assembly 62, and a diffuser case module 64.
  • the outer combustor wall assembly 60 and the inner combustor wall assembly 62 are spaced apart such that a combustion chamber 66 is defined therebetween.
  • the combustion chamber 66 is generally annular in shape to surround the engine central longitudinal axis A.
  • the outer combustor liner assembly 60 is spaced radially inward from an outer diffuser case 64A of the diffuser case module 64 to define an outer annular plenum 76.
  • the inner combustor liner assembly 62 is spaced radially outward from an inner diffuser case 64B of the diffuser case module 64 to define an inner annular plenum 78. It should be appreciated that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further appreciated that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.
  • the combustor wall assemblies 60, 62 contain the combustion products for direction toward the turbine section 28.
  • Each combustor wall assembly 60, 62 generally includes a respective support shell 68, 70 which supports one or more liner panels 72, 74 mounted thereto arranged to form a liner array.
  • the support shells 68, 70 may be manufactured by, for example, the hydroforming of a sheet metal alloy to provide the generally cylindrical outer shell 68 and inner shell 70.
  • Each of the liner panels 72, 74 may be generally rectilinear with a circumferential arc.
  • the liner panels 72, 74 may be manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material.
  • the liner array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that are circumferentially staggered to line the outer shell 68.
  • a multiple of forward liner panels 74A and a multiple of aft liner panels 74B are circumferentially staggered to line the inner shell 70.
  • the combustor 56 further includes a forward assembly 80 immediately downstream of the compressor section 24 to receive compressed airflow therefrom.
  • the forward assembly 80 generally includes a cowl 82, a bulkhead assembly 84, and a multiple of swirlers 90 (one shown). Each of the swirlers 90 is circumferentially aligned with one of a multiple of fuel nozzles 86 (one shown) and the respective hood ports 94 to project through the bulkhead assembly 84.
  • the bulkhead assembly 84 includes a bulkhead support shell 96 secured to the combustor walls 60, 62, and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96 around the swirler opening.
  • the bulkhead support shell 96 is generally annular and the multiple of circumferentially distributed bulkhead liner panels 98 are segmented, typically one to each fuel nozzle 86 and swirler 90.
  • the cowl 82 extends radially between, and is secured to, the forwardmost ends of the combustor walls 60, 62.
  • the cowl 82 includes a multiple of circumferentially distributed hood ports 94 that receive one of the respective multiple of fuel nozzles 86 and facilitates the direction of compressed air into the forward end of the combustion chamber 66 through a swirler opening 92.
  • Each fuel nozzle 86 may be secured to the diffuser case module 64 and project through one of the hood ports 94 and through the swirler opening 92 within the respective swirler 90.
  • the forward assembly 80 introduces core combustion air into the forward section of the combustion chamber 66 while the remainder enters the outer annular plenum 76 and the inner annular plenum 78.
  • the multiple of fuel nozzles 86 and adjacent structure generate a blended fuel-air mixture that supports stable combustion in the combustion chamber 66.
  • the outer and inner support shells 68, 70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT 54.
  • the NGVs 54A are static engine components which direct core airflow combustion gases onto the turbine blades of the first turbine rotor in the turbine section 28 to facilitate the conversion of pressure energy into kinetic energy.
  • the core airflow combustion gases are also accelerated by the NGVs 54A because of their convergent shape and are typically given a "spin” or a "swirl” in the direction of turbine rotor rotation.
  • the turbine rotor blades absorb this energy to drive the turbine rotor at high speed.
  • a multiple of studs 100 extend from each of the liner panels 72, 74 so as to permit a liner array (partially shown in Figure 4 ) of the liner panels 72, 74 to be mounted to their respective support shells 68, 70 with fasteners 102 such as nuts. That is, the studs 100 project rigidly from the liner panels 72, 74 to extend through the respective support shells 68, 70 and receive the fasteners 102 on a threaded section thereof ( Figure 5 ).
  • a multiple of cooling impingement passages 104 penetrate through the support shells 68, 70 to allow air from the respective annular plenums 76, 78 to enter cavities 106 formed in the combustor walls 60, 62 between the respective support shells 68, 70 and liner panels 72, 74.
  • the impingement passages 104 are generally normal to the surface of the liner panels 72, 74.
  • the air in the cavities 106 provides cold side impingement cooling of the liner panels 72, 74 that is generally defined herein as heat removal via internal convection.
  • a multiple of effusion passages 108 penetrate through each of the liner panels 72, 74.
  • the geometry of the passages e.g., diameter, shape, density, surface angle, incidence angle, etc., as well as the location of the passages with respect to the high temperature combustion flow also contributes to effusion cooling.
  • the effusion passages 108 allow the air to pass from the cavities 106 defined in part by a cold side 110 of the liner panels 72, 74 to a hot side 112 of the liner panels 72, 74 and thereby facilitate the formation of a thin, relatively cool, film of cooling air along the hot side 112.
  • each of the multiple of effusion passages 108 are typically 0.01-0.05 inches (0.254 - 1.27 mm) in diameter and define a surface angle of about 15-90 degrees with respect to the cold side 110 of the liner panels 72, 74.
  • the effusion passages 108 are generally more numerous than the impingement passages 104 and promote film cooling along the hot side 112 to sheath the liner panels 72, 74 ( Figure 6 ).
  • Film cooling as defined herein is the introduction of a relatively cooler air at one or more discrete locations along a surface exposed to a high temperature environment to protect that surface in the region of the air injection as well as downstream thereof.
  • impingement passages 104 and effusion passages 108 may be referred to as an Impingement Film Floatwall (IFF) assembly.
  • IFF Impingement Film Floatwall
  • a multiple of dilution passages 116 are located in the liner panels 72, 74 each along a common axis D.
  • the dilution passages 116 are located in a circumferential line W (shown partially in Figure 4 ).
  • the dilution passages 116 are illustrated in the disclosed non-limiting embodiment as within the aft liner panels 72B, 74B, the dilution passages may alternatively be located in the forward liner panels 72A, 72B or in a single liner panel which replaces the fore/aft liner panel array.
  • the dilution passages 116 although illustrated in the disclosed non-limiting embodiment as integrally formed in the liner panels, it should be appreciated that the dilution passages 116 may be separate components. Whether integrally formed or separate components, the dilution passages 116 may be referred to as grommets.
  • each of the forward liner panels 72A, 72B, and the aft liner panels 74A, 74B in the liner panel array includes a perimeter rail 120a, 120b formed by a forward circumferential rail 122a, 122b, an aft circumferential rail 124a, 124b, and axial rails 126Aa 126Ab, 126Ba, 126Bb, that interconnect the forward and aft circumferential rail 122a, 122b, 124a, 124b.
  • the perimeter rail 120 seals each liner panel with respect to the respective support shell 68, 70 to form the impingement cavity 106 therebetween.
  • the forward and aft circumferential rail 122a, 122b, 124a, 124b are located at relatively constant curvature shell interfaces while the axial rails 126Aa 126Ab, 126Ba, 126Bb, extend across an axial length of the respective support shell 68, 70 to complete the perimeter rail 120a, 120b that seals the forward liner panels 72A, 72B, and the aft liner panels 74A, 74B to the respective support shell 68, 70.
  • a multiple of studs 100 are located adjacent to the respective forward and aft circumferential rail 122a, 122b, 124a, 124b.
  • Each of the studs 100 may be at least partially surrounded by posts 130 to at least partially support the fastener 102 and provide a stand-off between each forward liner panels 72A, 72B, and the aft liner panels 74A, 74B and respective support shell 68, 70.
  • the dilution passages 116 are located downstream of the forward circumferential rail 122a, 122b in the aft liner panels 72B, 74B to quench the hot combustion gases within the combustion chamber 66 by direct supply of cooling air from the respective annular plenums 76, 78. That is, the dilution passages 116 pass air at the pressure outside the combustion chamber 66 directly into the combustion chamber 66.
  • the dilution passages 116 include at least one set of circumferentially alternating major dilution passages 116A and minor dilution passages 116B (also shown in Figure 5 ). That is, in some circumferentially offset locations, two major dilution passages 116A are separated by one minor dilution passages 116B. Here, every two major dilution passages 116A are separated by one minor dilution passages 116B but may still be considered "circumferentially alternating" as described herein.
  • either or both of the combustor liner support shells 68, 70 is formed with a furrow 150 to receive a rail 122 that extends from a respective liner panel 72, 74.
  • the furrow 150 may be formed at one or a multiple of locations in the combustor liner support shells 68, 70 to receive, a portion of the perimeter rail 120a, 120b, such as the forward and aft circumferential rail 122a, 122b, 124a, 124b at an interface between the forward liner panel 72A and the aft liner panel 74A; a single forward or aft circumferential rail 122a, 122b, 124a, 124b of the forward liner panel 72A or the aft liner panel 74A ( Figure 8 ); and/or an intermediate rail 126 ( Figure 9 ).
  • the furrow 150 includes a displaced surface 152 generally parallel to a combustor liner support shell inner surface 154 that faces the liner panels 72, 74, a first surface 156 that extends between the displaced surface 152 and the combustor liner support shell inner surface 154, and a second surface 158 that extends between the displaced surface 152 and the combustor liner support shell inner surface 154.
  • the combustor liner support shell inner surface 154 is the primary surface of the combustor liner support shell 68, 70. That is, the majority of the combustor liner support shell 68, 70 forms the combustor liner support shell inner surface 154 which faces the cold side 110 of the liner panels 72, 74.
  • the furrow 150 receives the respective aft circumferential rail 124a of the forward liner panel 72A and the forward circumferential rail 122a of the aft liner panels 74A.
  • the surfaces 156, 158 include interfaces 160, 162, 164, 166, between the displaced surface 152 and the combustor liner support shell inner surface 154 that may be formed as radiuses, chamfers, or other shapes in response to, for example, hydroforming of the combustor liner support shell 68, 70.
  • the surfaces 156, 158 may include impingement passages 104 that direct airflow directly onto the respective forward circumferential rail 122a and the aft circumferential rail 124a.
  • the respective forward circumferential rail 122a and the aft circumferential rail 124a that are received into the furrow 150 are generally twice the height of the standard rail height which spans the cavities 106. That is, the furrow 150 defines a depth generally equal to the height of the cavity 106 such that the rails which are received into the furrow 150 are correspondingly of twice the height.
  • the respective forward circumferential rail 122a and the aft circumferential rail 124a bottom out on the displaced surface 152 of the furrow 150 at assembly, such that an interference fit is achieved at operational temperatures.
  • the dilution passage 116 is defined by walls 160 that are received into a furrow 150A that, in this embodiment, is circular. That is, the walls 160 are essentially recessed into the combustor liner support shell 68, 70. It should be appreciated that various other liner panel interfaces including, for example, an igniter hole will also benefit herefrom.
  • the combustor panel and shell arrangement allows for cooling to increase panel durability.
  • the interface between the panel and shell also provides an arrangement that enables reduced leakage adjacent to the rails.
  • the furrows also serve as dirt and particle separators given the they create ribs about the exterior of the shell.

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  • Engineering & Computer Science (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Mechanical Engineering (AREA)
  • General Engineering & Computer Science (AREA)
  • Ceramic Engineering (AREA)
  • Turbine Rotor Nozzle Sealing (AREA)

Claims (9)

  1. Brennkammer (56) für ein Gasturbinentriebwerk (20), Folgendes umfassend:
    eine Brennkammerliner-Trägerschale (68, 70) mit einer darin gebildeten Furche (150);und
    eine Linerplatte (72A, 74A), die eine Schiene (124a) aufweist, wobei sich die Schiene (124a) in die Furche (150) erstreckt;
    gekennzeichnet durch:
    eine zweite Linerplatte (74A), die eine zweite Schiene (122a) aufweist, wobei sich die zweite Schiene (122a) in die Furche (150) erstreckt.
  2. Brennkammer (56) nach Anspruch 1, wobei die Schiene (124a) mindestens teilweise eine Fläche berührt, die durch die Furche (150) gebildet ist.
  3. Brennkammer (56) nach Anspruch 1 oder 2, wobei die Furche (150) mindestens einen dort hindurch verlaufenden Prallkühlungskanal (104) beinhaltet, wobei der Prallkühlungskanal (104) auf die Schiene gerichtet ist.
  4. Brennkammer (56) nach Anspruch 1, 2 oder 3, wobei die zweite Linerplatte (74A) eine hintere Linerplatte ist.
  5. Brennkammer (56) nach einem der vorstehenden Ansprüche, wobei die Furche (150) eine konkave Fläche bildet, die der Linerplatte (72A, 74A) zugewandt ist.
  6. Brennkammer (56) nach Anspruch 1, wobei die Linerplatte eine vordere Linerplatte (72A) ist, die an der Trägerplatte (68, 70) über mehrere Bolzen (100) montiert ist, die Schiene eine vordere Linerplattenschiene (124a) ist, die sich in die Furche (150) erstreckt, die zweite Linerplatte eine hintere Linerplatte (74A) ist, die an der Trägerschale (68, 70) über mehrere Bolzen (100) stromabwärts der vorderen Linerplatte (72A) montiert ist, und die zweite Schiene eine hintere Linerplattenschiene (122a) ist, die sich in die Furche (150) erstreckt.
  7. Brennkammer (56) nach Anspruch 6, wobei die Furche (150) mindestens einen dort hindurch verlaufenden Prallkühlungskanal (104) beinhaltet, wobei der Prallkühlungskanal (104) auf die Schiene (124a) gerichtet ist.
  8. Brennkammer (56) nach Anspruch 6 oder 7, wobei die Furche (150) eine versetzte Fläche (152), die parallel zu der Brennkammer-Trägerschalenfläche ist und der ersten und der zweiten Linerplatte (72A, 74A) zugewandt ist, eine erste Fläche (156), die sich zwischen der versetzten Fläche (152) und der Fläche erstreckt, und eine zweite Fläche (158), die sich zwischen der versetzen Fläche (152) und der Fläche erstreckt, beinhaltet.
  9. Brennkammer (56) nach Anspruch 8, wobei die erste Fläche (156) und die zweite Fläche (158) in Bezug auf die Fläche und die versetzte Fläche (152) angewinkelt sind.
EP18156681.1A 2017-02-14 2018-02-14 Brennkammer für eine gasturbine Active EP3361158B1 (de)

Applications Claiming Priority (1)

Application Number Priority Date Filing Date Title
US15/432,098 US10739001B2 (en) 2017-02-14 2017-02-14 Combustor liner panel shell interface for a gas turbine engine combustor

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EP3361158A1 EP3361158A1 (de) 2018-08-15
EP3361158B1 true EP3361158B1 (de) 2019-09-04

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US10823411B2 (en) 2017-02-23 2020-11-03 Raytheon Technologies Corporation Combustor liner panel end rail cooling enhancement features for a gas turbine engine combustor
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